The present invention generally relates to turbine vanes and blades and, more particularly, to high temperature turbine vanes and blades designed for high effectiveness cooling and ease of manufacture.
Gas turbine power plants are used as the primary propulsive power source for aircraft, in the forms of jet engines and turboprop engines, as auxiliary power sources for driving air compressors, hydraulic pumps, etc. on aircraft, and as stationary power supplies such as backup electrical generators for hospitals and the like. The same basic power generation principles apply for all of these types of gas turbine power plants. Compressed air is mixed with fuel and burned, and the expanding hot combustion gases are directed against stationary turbine vanes in the engine. The vanes turn the high velocity gas flow to impinge upon turbine blades mounted on a turbine disk or wheel that is free to rotate.
The force of the impinging gas causes the turbine disk to spin at high speed. Jet propulsion engines use this power to draw more air into the engine and then high velocity combustion gas is passed out the aft end of the gas turbine, creating forward thrust. Other engines use this power to turn a propeller or an electric generator.
The turbine vanes and blades lie at the heart of the power plant, and it is well established that, in most cases, they are one of the limiting factors in achieving improved power plant efficiency. In particular, because they are subjected to high heat and stress loadings as they are rotated and impacted by the hot gas, there is a continuing effort to identify improvements to the construction and/or design of turbine vanes and blades to achieve higher performance.
Modern aircraft jet engines have employed internal cooling of turbine vanes and blades to keep the vane and blade temperatures within design limits. Typically, the vanes and blades are cooled by air (typically bled from the engine's compressor) passing through longitudinally extending internal passages, with the air entering near the vane endwalls or blade root (the attached portion of the blade). Known turbine vane and blade cooling techniques include a cooling circuit consisting of series-connected longitudinally-oriented passages producing serpentine flow which increase cooling effectiveness by extending the length of the coolant flow path.
A plateau for high temperature turbine vanes and blades has slowed progress toward more efficient engines. A slowing of cooling effectiveness improvement has been reached wherein cooling air is fed to the inside of the turbine vane or blade to be exhausted through small passages over the vane or blade and through the trailing edge. A typical turbine vane or blade utilizing this prior art is shown in U.S. Pat. No. 5,813,835. The concept of a multi-walled turbine vane or blade has been discussed for many years with attempts proving to be extremely costly to fabricate.
U.S. Pat. No. 5,328,331 for Turbine Airfoil With Double Shell Outer Wall discloses a blade that is similar to the blade configuration of the present invention in that it does have an outer wall and an inner cooler wall, but the cooling scheme for this prior art blade differs significantly from the inventive blade. This prior art blade utilizes an impingement scheme that requires a plurality of impingement holes in the cool inner wall to distribute the cooling flow to the outer wall. In contrast, the present invention's vane or blade flow circuits do not require impingement holes or any cooling flow through the large center body core passage. In fact, it is not always desirable to have flow through the center body core cavity, because a no/low flow condition in the center body core means a very low heat transfer coefficient for the inner cool wall. This feature of the inventive vane and blade minimizes the thermal gradient between the inner and outer walls.
U.S. Pat. No. 5,813,835 for Air-Cooled Turbine Blade discloses multiple center cavities that are used for cooling. The inventive blade utilizes one large center cavity that does not require cooling air. Thus, the inventive structure is lighter because it does not have multiple ribs dividing it into multiple cavities. In addition, because the inventive blade does not require cooling air in the center body core, it can better tailor the thermal gradient between the outer hot walls and the inner cooler walls.
This prior art vane or blade does utilize multi-pass cooling passages on portions of the pressure and suction sides of the airfoil, yet several important differences relative to the inventive blade are noted. First, the forward portion of the '835 blade is cooled with conventional flow circuits that simultaneously cool the pressure and suction surfaces. This does not allow independent, optimized cooling for the pressure and suction sides in the forward region of the blade as does the inventive blade. The inventive blade utilizes a forward flowing pressure side circuit, which then is used to cool the leading edge cavity. Moreover, it utilizes an aft flowing suction side circuit that is also used to cool the tip of the blade and is then recycled to continue to cool the aft portions of the blade to maximize the thermal effectiveness of the blade. The multi-pass circuits disclosed in this prior art patent exit out film holes and do not continue to form the leading and trailing edge cooling circuits as does the inventive vane or blade. In addition, the pressure and central cavity cooling circuits in the '835 blade are not independent as are the inventive blade flow circuits. This is a feature of the present vane or blade invention making it producible as individually separate cores. This prior art disclosure makes no mention of special flow enhancements to the serpentine turns using turning vane and pin placement, nor does it mention any out of plane turning which the inventive blade aft bend utilizes. This prior art blade does not utilize a tip plenum cooling circuit nor does it recycle the tip cooling air. There is no mention of trip strips in the tip cooling region or any tip flag cooling enhancements. This prior art blade does use a trailing edge flow discharge, but there is no mention of special placement of pin fins upstream of the trailing edge teardrops for vorticity control and film cooling enhancement.
U.S. Pat. No. 5,626,462 for Double-Wall Airfoil discloses a multi-walled airfoil construction but its cooling configuration and manufacturing method and method of construction are very different from the inventive blade. This prior art blade requires an airfoil skin material that is deposited on the inner airfoil support wall to produce the cooling cavities. Unlike the inventive blade, which is integrally cast as a single piece to produce the cooling circuits, this prior art blade requires that the inner support structure be machined to create recessed grooves which can be made into cooling cavities later after the outer skin material is deposited. The '462 patent refers more to a method of construction of an airfoil structure than to a cooling configuration, which is described in vague generalities. This prior art blade cannot utilize cast pin fins (pedestals) or cast turning vanes in conjunction with pin fin placement for flow and heat transfer optimization in the flow channels as the inventive blade does, and it does not use tip cap cooling which gets recycled into the various cooling channels as does the inventive blade.